One example of a rocket engine is the RL10 rocket engine manufactured by the Pratt & Whitney Group of the assignee. The three major components of the engine are a turbopump, a combustion chamber and a rocket nozzle.
During operation of the rocket engine, the turbopump is used to supply a fuel and oxidant, such as liquid oxygen and liquid hydrogen, to the combustion chamber. The liquid oxygen and liquid hydrogen are expanded in the combustion chamber and burned to produce hot, pressurized rocket gases. The hot, pressurized gases are flowed at high velocities to the exhaust nozzle. The exhaust nozzle allows further expansion of the gases to increase the velocity of the gases before the gases exit the engine, thereby increasing the thrust of the rocket.
The exhaust nozzle is fabricated from thin walled tubes that are tapered and shaped to form the required nozzle contour. Liquid hydrogen fuel is flowed through these tubes to provide convective cooling to the tubes and regenerative heating to the hydrogen fuel. The convective cooling ensures that the temperature of the tubes is consistent with the temperature limits required for structural integrity of the nozzle.
Two important parameters of an exhaust nozzle are the expansion ratio of the nozzle at the exit and the pressure ratio of the nozzle under operative conditions. The pressure ratio is the pressure P.sub.c of the exhaust nozzle divided by the pressure P.sub.a of the ambient environment to which the rocket gases are discharged. The expansion ratio at any location of the nozzle is the area of the nozzle divided by the area of the throat of the nozzle.
The exhaust nozzle will flow "full" without separation of the boundary layer of the gases from wall of the nozzle if the pressure ratio is high enough. However, if the pressure ratio is too low for the amount of expansion required, then the flow will separate from the nozzle wall or casing which bounds the flowpath for the rocket gases.
This is caused in part by the static pressure in the wall region dropping well below ambient as the flow expands through a supersonic, overexpanded nozzle. As shown in FIG. 1, the boundary layer thickness increases with area ratio and the boundary layer has a portion which is subsonic. Ambient pressure is transmitted upstream through the subsonic portion of the boundary layer. Under the influence of this adverse pressure gradient, the subsonic portion of the boundary layer slows and thickens. The flow separates from the nozzle wall and induces a formation of shock waves which cause the primary overexpanded supersonic flow to adjust to the pressure of the inflowing ambient air.
For the three-dimensional-flow nozzles, the flow separation begins at a point on the nozzle wall rather than at a single area ratio plane normal to the centerline. This is because the pressure in the boundary layer for any one area ratio is subject to the random fluctuations that are characteristic of turbulent flow.
The ambient pressure of the in flowing air varies with altitude. At sea level, the pressure on the exterior of the nozzle and of the inflowing air is about fourteen and seven tenths pounds per square inch (14.7 psia). The total pressure on the interior of the nozzle, for example at a downstream location, might approach fifteen hundred pounds per square inch (1500 psi) with a static pressure of about fifteen pounds per square inch (15 psia). The pressure P.sub.b in the boundary layer might be in the order of two or three (2-3) pounds per square inch. The pressure differential may impose large side loads on the wall casing of the nozzle which surrounds the gas flow path.
Separation such as occurs when using a high area ratio exhaust nozzle at low pressure ratios may cause unstable flow separation, that is, the separation location may oscillate fore and aft with high frequency. This results in large fluctuating side loads. The fluctuating side loads are particularly troublesome in rockets because of the effect the side loads have on the trajectory of the rocket and because the side loads can cause destructive vibrations in the exhaust nozzle.
As the rocket driven vehicle travels from sea level to a high altitude above the earth's surface, (for example, 200,000 feet) the rocket nozzle will operate at pressure ratios (P.sub.c /P.sub.a) which vary dramatically from low pressure ratios at sea level to high pressure ratios at altitude. This results from the large variation in ambient pressure which decreases from about fourteen pounds per square inch (14.7 psia) at sea level to pressures as low as one-tenth of a pound per square inch (0.1 psia) at one hundred thousand feet (100,000 ft.) of altitude. Thus, it is difficult to avoid separation of flow in such a rocket nozzle.
FIG. 2 is an exemplary curve showing the relationship of the separation area ratio A.sub.r for a nozzle as a function of the pressure ratio at which the nozzle is operated. For a given pressure ratio of the rocket nozzle, separated flow will occur in the chamber at area ratios above the separation area ratio. For example, at a pressure ratio of two hundred the exhaust nozzle will flow "full" at an area ratio less than about 68 and will flow separated at an area ratio greater than 68 with overexpansion of the flow and unstable separation occurring at an axial location corresponding to the about the area ratio 62 of the nozzle. If separation were the only concern, the ideal nozzle would have an area ratio at the exit which is equal to the area ratio at which separation occurs.
One solution, then, is to employ an exhaust nozzle having a maximum area ratio which, for the operative pressure ratios, causes the nozzle to flow full from take-off until the rocket reaches its final altitude. However, this reduces the efficiency of the rocket at high altitudes because of the low velocities of the exiting flow. These low velocities are associated with the small expansion ratio of area used to avoid separation at low pressure ratios which occur at sea level take-off.
Creating a vehicle having several stages is another approach to tailoring the nozzle area ratio and pressure ratio to avoid separation. Each stage has a nozzle having area ratios designed for a limited range of pressure ratios. The nozzle would have a stage having low area ratio at sea level launch and another stage having a high area ratio at high altitude orbit. Multi-stage rocket engines are more complex than single stage rocket engine and increase the complexity of operation. However, such a rocket has the advantage of increased thrust as the rocket reaches orbit.
The above art notwithstanding, scientists and engineers have been working under the direction of Applicants' assignee to develop a rocket engine having an exhaust nozzle which would eliminate the requirement for staging and having a high expansion ratio (that is, greater than 150:1) for good thrust at altitude while not having unacceptably large unstable separation.